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AD-A244 669 4
NAVAL POSTGRADUATE SCHOOL L7Monterey, California
DTIct'DLA 'J I 9L CC'
JAN 3.199
THESISBaseline Vibration Measurements of Remotely Piloted
Helicopters for Higher Harmonic Control Research
by
Kevin M. Ransford
December, 1991Thesis Advisor: E. Roberts Wood
Approved for public release; distribution is unlimited
92-01581
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Baseline Vibration Measurements of Remotely Piloted Helicopters for Higher Harmonic Control Research (U )
12. PERSONAL AUTHOR(S) Ranslord, Kevin M.
13a TYPE OF REPORT 13b TIME COVERED 14. DATE OF REPORT (year, month, day) 15 PAGE COUNTMaster's Thesis From To December 1991 7116. SUPPLEMENTARY NOTATION
The views expressed in this thesis are those ofthe author and do not reflect the official policy or position of the Department of Defense or the U.S.Government.17 COSATI CODES 18. SUBJECT TERMS (continue on reverse if necessary and identify by block number)
FIELD GROUP SUBGROUP IIIIC, RPH, Vibrations, Flight Tests, Ground Tests
19 ABSTRACT (continue on reverse if necessary and identify by block number)
The Department of Aeronautics and Astronautics at the Naval Postgraduate School (NPSI is conducting a research program in the methods ofhigher harmonic control (HHC) for reduction of helicopter vibrations. The program at NPS uses remotely piloted helicopters (RPHI to studyHHC effects on vibration and blade load reduction. The scope of this masters thesis was to measure the baseline vibration profile of the RPH testvehicles prior to the installation of a HHC system. This goal was met by the development of a data instrumentation and recording system and byconducting a ground and flight test program for the RPH test vehicles. From the results of these tests it was concluded that: a) the datainstrumentation and recording system was of sufficient sensitivity to detect vibrations experienced within the RPH airframe, and, bi the RPHexhibited a vibration profile similar to that of a full scale helicopter. It is recommendcd that a HHC system be designed, fabricated, and installedon the Pacific RPV Co. 'Bruiser' so that the effects of HHC on helicopter performance may be evaluated.
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DO FORM 1473,84 MAR 83 APRedition may be used u.ntil exhausted SECcuF!- ,LrS;:,CATION OF THIS PAGEAll other editions are obsolete U nclassified
Approved for public release; distribution is unlimited.
Baseline Vibration Measurements ofRemotely Piloted Helicopters For Higher Harmonic Control Research
by
Kevin M. RansfordB.S., Parks College of St. Louis University, 1985
Rotary Wing Aircraft Test DirectorateNaval Air Test Center
Submitted in partial fulfillmentof the requirements for the degree of
MASTER OF SCIENCE IN ENGINEERING SCIENCE (AERONAUTICS)
from the
NAVAL POSTGRADUATE SCHOOLDECEMBER, 1991
Author: 14, /7/ Kevig .Rasford-
Approved by: C 2 . 0QE. Roberts Wood, Thesis Advisor
Richard M. Howard, Second Reader
Daniel J. Collins, hairmanDepartment of
Aeronautics and Astronautics
ii
ABSTRACT
The Department of Aeronautics and Astronautics at the
Naval Postgraduate School (NPS) is conducting a research
program in methods of higner harmonic control (HHC) for
reduction of helicopter vibrations. The program at NPS uses
remotely piloted helicopters (RPH) to study HHC effects on
vibration and blade load reduction. The scope of this
master's thesis was to measure the baseline vibration profile
of the RPH test vehicles prior to the installation of a HHC
system. This goal was met by the development of a data
instrumentation and recording system and by conducting a
ground and flight test program for the RPH test vehicles.
From the results of these tests it was concluded that: a) the
data instrumentation and recording system was of sufficient
sensitivity to detect vibrations experienced within the RPH
airframe; and b) the RPH exhibited a vibration profile
similar to that of a full scale helicopter. It is recommended
that a HHC system be designed, fabricated, and installed on
the RPH so that the effects of HHC on helicopter performance
may be evaluated. Acoess or
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TABLE OF CONTENTS
I. INTRODUCTION ........................................ 1
II. FUNDAMENTALS OF HIGHER HARMONIC CONTROL ............... 4
III. DESCRIPTION OF RPH TEST VEHICLES .................... 6
A. The GMP Legend ................................. 7
B. The Pacific RPV Co. Bruiser .................. 10
IV. DESCRIPTION OF TEST INSTRUMENTATION ................... 14
A. Scientific-Atlanta SD380 Signal Analyzer ..... 15
B. PCB Corp. Series 302B03 Accelerometer ......... 16
C. PCB Corp. Model 480B Battery Power Unit ...... 17
D. The RPH Data Acquisition System ................ 18
V. METHOD OF TESTS ..................................... 20
A. GMP Legend Tests ............................. 20
1. Fuselage Vibration Measurements ........... 21
2. Tail Boom Measurements ..................... 24
3. Main Rotor Imbalance Measurements ......... 25
B. Pacific RPV Co. Bruiser Tests .................. 27
1. Ground Tests ............................. 27
VI. RESULTS AND DISCUSSION .............................. 30
A . General ..................................... 31
B. GMP Legend Measurements ..................... 31
1. Fuselage Measurements ..................... 31
2. Tail Boom Measurements .................... 32
iv
3. Main Rotor Mass Imbalance Measurements.. 34
C. Pacific RPV Co. Bruiser Measurements ......... 36
VII. CONCLUSIONS AND RECOMMENDATIONS .................... 38
A. Conclusions .................................. 38
1. General .................................. 38
2. The GMP Legend ........................... 38
3. The Pacific RPV Co. Bruiser ................ 39
B. Recommendations .............................. 40
1. General .................................. 40
2. Specific ................................. 40
APPENDIX A: MODEL 480B BATTERY SUPPLY UNIT/ MODEL 302B03
ACCELEROMETER DATA SHEETS ..................... 42
APPENDIX B: MODEL 302B03 ACCELEROMETER CALIBRATION DATA... 44
APPENDIX C: GMP LEGEND FUSELAGE MEASUREMENT DATA ........... 46
APPENDIX D: GMP LEGEND TAIL BOOM MEASUREMENT DATA .......... 49
APPENDIX E: MAIN ROTOR MASS IMBALANCE MEASUREMENTS ......... 51
APPENDIX F: PACIFIC RPV CO. BRUISER MEASUREMENT DATA ...... 54
LIST OF REFERENCES ........................................ 62
INITIAL DISTRIBUTION LIST ................................. 64
v
I. INTRODUCTION
Analysis and control of helicopter vibrations is an area
of vital concern to the helicopter designer, operator, and
mechanic. Excessive vibration levels are detrimental to the
helicopter airframe and its components, and can also lead to
the early onset of flight crew fatigue. Therefore, reduction
of vibration levels to a minimum acceptable value remains of
prime importance. Fortunately, continued research into
methods for the reduction of vibrations has produced a gradual
reduction in vibration levels [Figure 1, (Ref I)].
U bO
0
050o40 MIL H450IA SMIFICATION
z030 0 AAHIUTTAS SPECIFICATION
03 REVISD AAHUTTAS SPECIFICATION
z o
z 010 --[
Z 0 0 G D E S IA 1 0- - - - -. .00
195S 190 196S 1970 1975 1960 19s
YEARS
Figure 1. The Trend of Helicopter Vibration Levels
Methods for reduction of helicopter vibrations include
both active and passive measures. Passive measures include
mechanical vibration absorbers and isolators. Active measures
include individual blade control (IBC) [Ref 2.] and higher
harmonic control (HHC). As shown in Figure 1, the level of
vibration reduction that may be realized by passive methods
has reached its maximum potential. Thus, an active means of
reducing this level of vibration now appears to provide the
best possible solution.
Research in the application of HHC for vibration control
has been conducted since the early 1960's to investigate its
effect on both vibration and blade-load reduction. References
3 through 9 summarize the results of this research. Reference
3 is noteworthy in that it was the only actual flight test
program using the HHC concept until the 1980's.
Reference 1 reported on development and flight tests of
HHC by Hughes Helicopters, Inc. (now McDonnell Douglas
Helicopter Co.) under a joint U.S. Army/National Aeronautics
and Space Administration (NASA) program using an OH-6A
helicopter during 1982-83. Flight test results presented in
this study showed that the HHC system provided a significant
reduction in helicopter vibrations without penalties in blade
loads or aircraft performance. Further analysis of these HHC
flight test results showed a possible performance increase
for the helicopter using HHC. This increase was
characterized by a twenty per-cent (20%) reduction in required
main rotor shaft torque [Ref. 101. However, the scope of the
HHC test program [Ref. 1] and funding limitations prevented
additional flight testing and evaluation of the effects of HHC
on helicopter performance.
2
Thus, important questions regarding the effects of HHC on
helicopter performance remain unanswered. How will it change
power available? That is, does minimum power occur at the
same phase as minimum vibration? At what airspeeds will
performance show the greatest increase (or decrease)? What is
the effect of HHC on fuel consumption? It is with these
questions in mind that the HHC research program at the Naval
Postgraduate School (NPS) was established.
This NPS HHC program has been conducted by master's level
thesis students with the intent of developing an active flight
test program. This flight test program utilizes remotely
piloted helicopters (RPH) as the primary flight test
vehicles. Previous research efforts at NPS [Ref's. 11
through 15] focused primarily on the procurement,
construction, and calibration of the RPH, and on the
establishment of a vertical flight test capability at the NPS
Unmanned Air Vehicle (UAV) Flight Research Laboratory. The
scope of this masters thesis is to conduct an initial ground
and flight test investigation of the RPH test vehicles
vibration spectrum, prior to modification of the RPH to the
HHC system.
3
II. FUNDAMENTALS OF HIGHER HARMONIC CONTROL
Higher harmonic control of helicopter vibrations is a
method by which an active control system is used to control
and reduce the level of helicopter vibrations. The vibrations
of greatest magnitude result from aerodynamic loads acting
on the main rotor system. Dynamic coupling of these loads
between the rotor system and the airframe results in a
measurable vibration profile for the aircraft. Other sources
of vibrations include the engine and tail rotor, which provide
excitation at a reduced level compared with those imparted
to the airframe by the main rotor system.
Reduction of the level of vibrations is accomplished using
HHC by measuring the vibrations which are transmitted to the
airframe, and then imparting a motion back to the main rotor
system (blade pitch) at a frequency which will reduce or
cancel the dominant vibration frequency. For a four-bladed
helicopter, this method is normally applied as foilows: Tri-
axial accelerometers located at the pilot's seat in the
airframe sense vertical, lateral, and longitudinal
vibrations. These vibrations are converted to electrical
signals and are then routed to an electronic processing unit.
The processing unit analyzes the accelerometer signals and
determines the level of blade feathering required to cancel
the vibration. A superimposed 4/rev. swashplate motion is
4
applied to collective and cyclic control inputs. Perturbing
the stationary swashplate at 4/rev. both collectively and in
pitch and roll results in third, fourth, and fifth harmonic
blade feathering (typically less than one degree) in the
rotating system. This feathering action results in reduced
excitation from the rotcr system, which acts as a filter and
transmits the corresponding lower vibration levels to the
airframe. Pitch, roll, and collective motion of the
stationary swashplate at this 4/rev. frequency is provided by
electro-hydraulic or electronic high frequency servo-
actuators. Depending on system design, HHC may be operated in
the open- or closed-loop mode. [Ref. 1] However, HHC
research at NPS will be conducted initially in the oper-loop
format.
5
III. DESCRIPTION OF RPH TEST VEHICLES
Development of the ground and flight test program at NPS
(using the RPH test vehicles) has been based upon a
systematic, gradual build-up of instrumentation, testing
milestones, and pilot proficiency prior to installation and
flight test of a RPH with HHC. This test program was
designed so that initial instrumentation development, data
collection techniques, and pilot training could all be
conducted on a RPH that was easy and inexpensive to fly,
modify, and maintain. Pilot training was considered of
paramount importance, as it placed a severe limitation on data
collection and analysis in previous thesis studies [Ref. 151
These considerations led to the procurement of two RPH test
vehicles. RPH test vehicles were chosen for this research
effort because th( offer a means of investigating the
properties of HHC without the cost of a full scale flight test
program. In addition, the RPH uses control systems and
main/tail rotor blade airfoil sections similar to a full scale
helicopter. Thus, it is expected that the RPH test vehicles
would yield research results which may be useful to the full
cale helicopter designer.
The first of these test vehicles, the GMP Legend, is a
two-bladed aircraft with a stabilizing flybar, a two-bladed
tail rotor, and a .61 cubic inch engine. The second test
6
vehicle, the Pacific RPV Bruiser is a four-bladed aircraft
with a four-bladed tail rotor, and a four horsepower engine.
For this test program, the Legend was used as the
instrumentation development and pilot training aircraft, and
the Bruiser was used for initial ground-based vibration
measurement studies prior to the installation of the HHC
system. A primary factor in the selection of the Bruiser for
HHC work was that it had a four-bladed rotor. Since (n-i), n,
and (n+l) per rev. frequencies are involved, the (n-i)
component for a two-bladed rotor (i/rev.) couples into the
output of the primary control system, which is undesirable.
This problem is eliminated with the four-bladed rotor system.
These aircraft are described more fully in following
paragraphs.
A. THE GMP LEGEND
The GMP Legend, as noted previously, is a two-bladed
flybar stabilized RPH. This aircraft was chosen as the
initial development aircraft because it allowed the test team
to develop test methods, instrumentation, and data collection
techniques (with the added capability of being an excellent
training aircraft for the novice RPH pilot) prior to operating
the more difficult-to-fly Bruiser. The ease in maintenance
and the availability of spare parts were also fundamental
considerations in the choice of the Legend for the initial
stages of the test program. The Legend is presented Figure 2.
7
Control of the Legend is accomplished using a Futaba Corp.
Model F995HP nine channel transmit.er, which operates using a
pulse code modulation 'PCM) transmission technique. A radio
receiver, rate gyros, and control servos mounted on the RPH,
complete the system. An added advantage of this transmitter
system is that it is similar to the system used to control the
Bruiser RPH. Table I provides addi-ional specifications of
the GMP Legend.
Figure 2. The GP Legend
8
TABLE I. LEGEND SPECIFICATIONS
Main Rotor Diameter 58"Number of Rotor Blades 2Tail Rotor Diameter 10"Length 51"1Weight 8.8 lbsPayload 10 lbsEngine .61 cu. in.Operating RPM (Main Rotor) 1500 - 1700 RPMEngine to Main RotorGearing Ratio 8.6:1Main Rotor to Tail RotorGearing Ratio 4.75:1
Engine, main rotor, and tail rotor excitation frequencies
and gearing ratios provide key information for the vibration
researcher because they characterize the primary vibration
frequencies transmitted to the airframe. These are the
frequencies of interest for HHC, so that it can be effective
in eliminating or reducing vibration excitation at these
frequencies. For a RPH like the two-bladed Legend, the
troublesome frequencies will occur at main rotor frequencies
of 1/rev., 2/rev., 4/rev., and harmonics thereof. Engine and
tail rotor primary frequencies and harmonics will also be
present. It should be noted that these frequencies will be
valid for only one main rotor operating RPM. In other words,
a change in operating RPM will result in a change in the
primary frequencies and their harmonics. Table II provides a
list of characteristic frequencies for operation at 1502 RPM.
9
TABLE II. GMP LEGEND CHARACTERISTIC FREQUENCIES(MAIN ROTOR OPERATING 1502 RPM)
Main Rotor n/rev. Frequency1/rev. 25 Hz.2/rev. 50 Hz.4/rev. 100 Hz.6/rev. 150 Hz.8/rev. 200 Hz.
Tail Rotor n/rev. Frequency
1/rev. 118 Hz.2/rev. 237 Hz.4/rev. 475 Hz.
Enqine n/rev. Frequency
1/rev. 215 Hz.2/rev. 460 Hz.
B. THE PACIFIC RPV CO. BRUISER
The Pacific RPV Co. Bruiser is a four-bladed helicopter
that was designed and built by the Pacific RPV Co., of
Startup, Washington. This RPV was procured for use as the HHC
test vehicle. Original design of this helicopter was for the
purposes of aerial spraying, photography, and electronic
countermeasures (ECM). The Bruiser is equipped with a four-
bladed, fully-articulated rotor head, with a four-bladed tail
rotor. Like the GMP Legend, the Bruiser utilizes a nine-
channel PCM transceiver system (using a Futaba Corp. FP-9VHP
transmitter) for operation of all flight control surfaces
and engine throttle movements. Electronic servos, controlled
by the airborne receiver (a Futaba Corp. FP-Rl29DP), operate
tail rotor, collective, and collective/cyclic pitch
10
mechanisms. The Bruiser is powered by a Super Tartan T77i
two-cylinder, two-cycle, air-cooled engine. This engine will
produce a maximum four brake horsepower (BHP) at 8,800 RPM,
but is typically operated at approximately 7,000 RPM. The
Bruiser is depicted in Figure 3. Additional specifications of
the Bruiser are contained in Table III.
.. j.
Figure 3. The Pacific RPV Co. Bruiser
TABLE III. THE PACIFIC RPV CO. BRUISER
Max. Forward Speed 50 KTSNumber of Rotor Blades 4Rotor Diameter 78 in.Tail Rotor Diameter 13 in.Empty Weight 18 lbs.Gross Weight 40 lbs.Main Rotor NormalOperating RPM 1175 RPMEngine to Main RotorGearing Ratio 8:1Main Rotor to Tail RotorGearing Ratio 6:1
11
As with the GMP Legend, engine, main rotor, and tail rotor
gearing ratios will be used to identify primary frequencies
and harmonics of the rotating components when the Bruiser is
operated at a specific rotor RPM. Table IV lists the
characteristic values of these frequencies for operation at
1175 RPM, which is normal operating speed.
TABLE IV. PACIFIC RPV CO. BRUISER CHARACTERISTIC FREQUENCIES(MAIN ROTOR OPERATING 1175 RPM)
Main Rotor n/rev. Frequency
1/rev. 19 Hz.2/rev. 39 Hz.4/rev. 78 Hz.6/rev. 117 Hz.8/rev. 156 Hz.
Tail Rotor n/rev. Frequency
1/rev. 117 Hz.2/rev. 234 Hz.
Enqine n/rev. Frequency
1/rev. 156 Hz.2/rev. 313 Hz.
12
It should be noted that main rotor 6/rev. and tail rotor
1/rev, are the same frequency of 117 Hz. This condition also
exists for main rotor 8/rev. and engine 1/rev. While this
condition will not result in unusual vibrations or operation
of the RPH, it will make contributions of the individual
components of the tail rotor 1/rev. and engine 1/rev.
vibration amplitudes much more difficult to separate and
analyze.
13
IV. DESCRIPTION OF TEST INSTRUMENTATION
Analysis and recording of helicopter vibrations on a full
scale helicopter is normally performed using selected
accelerometers and monitoring equipment. This equipment may
include an oscilloscope, a spectrum-type signal analyzer, or
a direct plotting device such as the Chadwick Helmuth Corp.
Vibration Test Set. Each of these equipments was considered
during development of a data measuring and recording system
for the RPH test vehicles. As with a full scale test program,
size, weight, and power constraints were primary factors in
considering what system or combinations of systems would best
serve the test program needs. In addition, the system chosen
must be sensitive enough to detect vibrations at various
frequencies from 1 to 1000 Hz., at amplitudes much less than
those encountered with a full scale helicopter. Review of the
available instrumentation options led to the choice of a
spectrum analyzer which could both store data for later
retrieval and which would interface with a plotter/printer
without additional hardware. A tri-axial mount with highly
sensitive accelerometers was procured for mounting on the RPH.
Individual components and a system description is provided in
greater detail in the following paragraphs.
14
A. THE SCIENTIFIC-ATLANTA SD380 SIGNAL ANALYZER
The Scientific-Atlanta SD380 Signal Analyzer is a four-
channel spectrum analyzer that is designed to provide
frequency domain display of various electronic signals. Up to
four channels of input data may be displayed in either a
logarithmic, Linear, or waterfall format. 'irs-r tracking,
statistical parameters, and :ext features allow -or complete
diata analysis without additional microcomputer cperations.
Display data may be stored on a 3.5 in. floppy disk via the
SD380's disk drive, and, plotting/printing operations may be
performed via the installed IEEE 488 data bus. These latter
two attributes were considered to be the SD330's most
important features for this research program. This unit is
cransportable via a common whe d laboratory cart, and is
powered by 115 volts AC. The SD380 is presented in Figure 4.
Figure 4. The SD380 Signal Analyzer
15
B. THE PCB CORP. SERIES 302B03 ACCELEROMETER
Measurement of the vibration levels during RPH operation
was provided by PCB Corp. Series 302B03 piezoelectric
accelerometers. These accelerometers are approximately 1.30
inches in length and .48 inches in diameter. This
accelerometer maybe attached to the test object by means of an
beryllium-copper stud or mounting wax. This particular model
has a weight of approximately 25 grams and operates on +24 to
27 volts DC. The highly sensitive response of this
accelerometer (300 millivolts/G) made it ideal for RPH
testing. This accelerometer is presented Figure 5.
It, ......
03. j S.,I
Figure 5. The PCB Corp. Series 302B03 Accelerometer
16
C. TE PCB CORP. MODEL 480B BATTERY POWER UNIT
The Series 302B03 accelerometer is powered using the PCB
Corp. Model 480B Battery Power Unit. This power unit is
designed to power low-impedance piezoelectric transducers such
as the Series 302B03. This unit also contains built-in
coupling circuitry to allow for direct connection to
oscilloscopes, recorders, and signal analyzers. Power for the
Model 480B power unit is provided by three 9 volt batteries
mounted internal to the power unit chassis. The power unit is
connected to the accelerometers using a miniature 25 ft.
coaxial cable. Connection from the Model 480B power unit to
the display device is by a similar coaxial cable, which
terminates using a standard BNC connector at the spectrum
analyzer input. This power supply is presented Figure 6.
Additional details of the Series 302B03 accelerometer and
Model 480B power unit are provided Appendix A.
MODIL 48 AOWC i UN,,
Figure 6. The PCB Corp. Model 480B Battery Power Unit
17
D. THE RPH DATA ACQUISITION SYSTEM
Design of the RPH data acquisition system was based on the
requirement for the system to operate both while the RPH was
on the ground and in a hover. This requirement was met by use
of the 25 ft. coaxial cables described in Section C. These
cables were long enough to allow the RPH to operate on the
ground and hover out of ground effect (HOGE). This capability
allowed for flight in the parking lot behind the NPS UAV
Laboratory. Operation of the RPH behind the UAV Laboratory
eliminated the need for portable power generators for the
signal analyzer, and had the additional advantage of reducing
the effects of wind upon test results.
The data acquisition system was designed to operate as
follows: RPH mounted Series 302B03 accelerometers detected
vibrations transferred to the airframe from the engine, main
rotor, and tail rotor. These vibrations were converted to
electrical signals within the accelerometer, which is powered
by the Model 480B power unit. These electrical signals were
transmitted via the coaxial cable to the Model 480B power
unit. The power unit contains a coupling capacitor which
allows for transfer of the accelerometer signal through the
power unit to the spectrum analyzer using another coaxial
cable without amplification or attenuation of the
accelerometer signal. Display of the accelerometer signal is
provided by the SD380 spectrum analyzer. This display was
then recorded on the 3.5 in. floppy disk installed in the
18
SD380's internal disk drive. Data displays stored on the disk
drive could then be plotted using a Hewlett-Packard Model
7475A Plotter via the IEEE 488 data bus. This data
acquisition system is depicted in Figure 7.
HP 7475APLOTTER
SD380 _ _
SPECTRUMANALYZER MODEL 480B SERIES 302B03
POWER UNITS ACCELEROMETERS
Figure 7. The 1PH Data Acquisition System
19
V. METHOD OF TESTS
Development of a test program to measure and record the
vibration levels experienced by the RPH test vehicles was
arranged so that a gradual build-up of test complexity and
pilot experience would yield sufficient data without undue
risk to equipment. For this reason, a test schedule was
developed using both the GMP Legend and the Pacific RPV Co.
Bruiser. GMP Legend operations would allow for refinement and
proof-of-concept testing for the RPH data acquisition system
while also allowing the RPH pilot to gain experience with
Legend handling qualities. Following proof-of-concept
testing/pilot training with the Legend, ground and flight
tests of the Bruiser would be performed. As noted in previous
thesis efforts [References 14 and 15] the lack of an
experienced RPH pilot was the single greatest obstacle to
successful RPH flight testing of HHC. Use of the Legend for
proficiency flying was and will continue to be an asset to the
HHC program at NPS.
A. GMP LEGEND TESTS
Build-up efforts using the Legend were initiated in order
to answer the following questions: 1) Will the data collection
system be safe to operate? 2) Does the system have adequate
sensitivity for the level of vibrations that may be
20
encountered? 3) What are the optimum locations for the
accelerometers? 4) Will the researcher be able to distinguish
the various contributors to the vibration levels? 5) And
finally, does the Legend exhibit vibration traits similar to
those for a full scale helicopter? These questions were
addressed in subsequent test operations.
In order to investigate the Legend vibration profile, the
accelerometers were mounted at either a fuselage location or
on the tail boom, forward of the tail rotor assembly. During
all test events, the Legend was established in a hover out of
ground effect. As with a full scale aircraft, HOGE is a
distance of at least one rotor diameter between the bottom of
the fuselage and the ground. For the Legend, this distance
was approximately 58 inches.
In order to ensure data accuracy, the accelerometers
utilized for these tests were calibrated by the manufacturer
prior to the beginning of tests. Calibration results may be
found in Appendix B.
1. GMP Legend Fuselage Vibration Measurements
Measurements of vibration levels at the Legend
fuselage were obtained by mounting two 302B03 accelerometers
on a tri-axial block and attaching the tri-axial block to the
underside of the aft skid support, along the centerline of the
aircraft, approximately two inches aft of the gross weight
center-of-gravity (CG). The tri-axial block was secured to
21
the skid with a #10-32 screw. The accelerometers were mounted
on the tri-axial block so that RPH vertical and lateral
accelerations could be detected. Coaxial cables were attached
to the accelerometers and secured to the skid using plastic
-tie-wraps. The coaxial cable was then routed to the Model
480B power unit, and then to the SD380 spectrum analyzer. The
Legend was then established in a hover. Analyzer displays of
vibration levels were observed in a logarithmic vertical
scale (vibration amplitude) using a frequency scale
(horizontal axis) of up to 1000 Hz. Rotor RPM was measured
using a Skytach Corp. strobe tachometer. Establishment of
rotor RPM allowed for the rapid calculation of the 1/rev.
frequency (main rotor). This frequency was determined by the
following equation:
Main Rotor 1/Rev=(Main Rotor RPM/60 Hz.)
Once the main rotor 1/rev. frequency is determined, tail
rotor, engine, and associated harmonics can be noted for the
given main rotor operating RPM. Figure 8 depicts the Legend
with accelerometers mounted for fuselage measurements. Flight
testing of the Legend's fuselage vibration level is depicted
in Figure 9.
22
- 'ACC ELEROMETERS r <
Figure 8. GMP Legend With Fuselage Accelerometer Mount
, .. .:. .. ...... ., < ,
Figure 9. GMP Legend Fuselage Vibration Measurements Testing
23
2. GMP Legend Tail Boom Measurements
Vibration lev--is on the tail boom were measured using
accelerometers mounted on the tri-axial block in a manner
similar to that utilized for fuselage measurements. For
this test, the block was secured to the tail boom 27 inches
aft of the main rotoi shaft, and 24.5 inches aft of the gross
weight CG. A. aluminum strap and block assembly secured the
tri-axial block to the tail boom, and tie-wraps were used to
fasten the coaxial cables to the tail boom and skid
assemblies. The Legend was established in a hover identical
to the rrethiod used for fuselage measurements, and vibr?.:ion
levels were observed and recorded with the SD380 spectrum
analyzer. Figure 0 depicts the Legend with the
accelerometers mountec n the tail boom. Flight testing of
tail boom vibration levels is depicted Figure 11.
ACCELEROMETERS
Figure 10. GMP Legend With Accelerometers Mounted On Tail Boom
24
Figure 11. GMP Legend Tail Boom Vibration Measurements
Testing
3. GMP Legend Main Rotor Imbalance Measurements
Measurements of the effects of a main rotor system
unbalance were performed to: 1) Make a qualitative analysis of
the data acquisition system sensitivity; and 2) Examine the
change in vibration levels experienced by the Legend with a
main rotor unbalance, in particular at the main rotor 1/rev
frequency. A rotor system unbalance was created by sliding
one of the two main rotor flybar balance masses in toward the
main rotor mast a distance 1.75 inches inboard from the
normal balance mass position. This mass, which weighs
approximately .5 ounces, was secured in the test position on
the flybar by means of a setscrew. For this test, the tri-
axial block with accelerometers was mounted on the tail boom
25
in a manner identical that used for tail boom vibration
tests. The aircraft was then established in a hover.
Vibration levels were monitored and re. corded using the SD380
spectrum analyzer. This est was repeated with the balance
mass at a position 3.5 inches inboard fr-- the normal balance
mass position. Figures 12 and 13 show the Legend main rotor
system flybar balance weight at the 1.75 and 3.5 inch test
-ositions.
BALANCING MASS
Figure 12. GNP Legend With Balance Weight 1.75 Inches Inboard
26
BALANCING MASS
Figure 13. GMP Legend With Balance Weight 3.5 Inches Inboard
B. PACIFIC RPV CO. BRUISER TESTS
1. Ground Tests
Test instrumentation and data collection methods
utilized for the Legend flight testing were readily adapted
for use on the Bruiser. An adaptor block machined from
aluminum bar stock allowed for the attachment of the tri-
axial block directly to the Bruiser's fuselage aft skid mount.
As with the Legend, two Series 302B03 accelerometers were
mounted on the tri-axial block for detection of vertical and
lateral accelerations. Coaxial cable routing and spectrum
analyzer initialization was identical to the methods used
27
during Legend flight testing. Figure 14 depicts the Bruiser
with acceleroreters mounted for fuselage measurements.
Ground tests were ccnducted by operating the engine and
rotor systems while the Bruiser remained on its skid
structure. Main rotor RPM -as measured using the Skytach
tachometer. Vibration levels, characte .stic frequencies, and
harmonics of rotatin omponents were oL served using the SD380
spectrum analyzer. . Drage of these results using the SD-
380's 3.5 inch disk drive assisted in data reduction and
analysis efforts.
ACCELEROTS
Figure 14. Pacific RPV Co. Bruiser With FuselageAccelerometer Mount
28
This test program was developed with the intent of
measuring the Bruiser's ground and "in-flight" vibration
levels. However, failure of the engine/main rotor drive
assembly clutching mechanism at the conclusion of ground tests
prevented flight measurements from being accomplished in time
to meet the established schedule.
29
VI. RESULTS AND DISCUSSION
A. GENERAL
All data collected during Legend and Bruiser testing was
stored in individual data files using the SD380 3.5 in. floppy
disk drive. Reduction and analysis of this data was
accompLished by recalling the respective data file to the
SD380 spectrum analyzer cathode ray tube (CRT) display. The
SD380 cursor function was then used to identify both frequency
and amplitude of the displayed acceleration peaks. The
scaling of the vertical axis (amplitude) was calibrated in a
logarithmic scale. This was the scale selected for the data
analysis reported in this thesis. If desired, a linear scale
could also be selected for the vertical axis when using the
SD380 spectrum analyzer. For these tests, the basic unit of
measurement established on the logarithmic vertical scale was
the user-created Engineering Unit (EU) sub-function. This EU
sub-function was initialized so that:
1 Engineering Unit = 1 "G"
The horizontal axis was displayed in units of frequency on a
linear scale. Measurement of the main rotor RPM for the
Legend and Bruiser permitted rapid calculation of the various
characteristic frequencies and their respective multiples for
the respective RPH. The CRT data display was transferred to
30
hard copy (paper) plots using the HP 7475A plotter. These
plots are presented in Appendices C, D, E, and F.
B. GMP LEGEND MEASUREMENTS
1. Fuselage Measurements
Combined vertical and lateral vibration results are
presented in Appendix C, Figure 1. From these plots, it was
noted that engine 1/rev. (1P) was the, dominant vibration
component in both the lateral and vertical axes. Main rotor
1/rev. or 1P1 was difficult to identify using the combined
lateral/vertical scale. It was observed during these tests
that the main rotor, tail rotor, and their associated
harmonics were difficult to identify if the RPH was still in
ground effect (IGE) (when the data was stored to the 3.5 in.
disk). This may explain the lack of useful data from the
combined lateral/vertical plot. Figure 2 is an expanded view
of the lateral vibration measurements. Engine 1P is most
evident at a value of .305 EU's (.305 G) . Main and tail rotor
frequency amplitudes remain at or below the noise floor, and
cannot be identified. Engine 1P, 2P, and main rotor 1P, 2P,
and 4P frequencies are clearly identified as vertical
accelerations in Figure 3. Maximum amplitude of main rotor 1P
was .07 G. Main rotor 2P appears to be a slightly greater
I A standard convention that has been adopted forconvenience in helicopter dynamics is to abbreviate thestatements I/rev., 2/rev.,...etc. by 1P, 2P,...etc. Thisnotation has been adopted in this thesis.
31
value than main rotor 1P, as would be expected. Noteworthy in
these fuselage measurements is the fact that the engine 1P is
of a greater value in the lateral axis (.305 G) than in the
vertical axis (approx. .10 G). In addition, main rotor 1P
and associated harmonics were observed when measured in the
vertical axis, but did not appear in lateral measurements.
Tail rotor 1P could not be identified in either the lateral or
vertical axes using the fuselage accelerometer mount.
Finally, engine 1P is the frequency of greatest amplitude as
compared with all other characteristic frequencies. This is
not typical of a full scale helicopter, where for a 2-bladed
design the main rotor 2P would be the dominant frequency, with
engine and drive train components of lesser magnitude.
2. Tail Boom Measurements
Vibration measurements at the Legend tail boom
accelerometer location are presented in Appendix D. Figure 1
of Appendix D is a combined lateral/vertical plot which again
shows the predominance of the engine 1P frequency in both
axes. Main rotor 1P and 2P frequencies (27.5 and 55 Hz,
respectively) are clearly defined when measured in the lateral
axis, but are not visible in the vertical axis. This is in
contrast to fuselage measurements, in which the main rotor 1P
and 2P characteristic frequencies were of greatest amplitude
when measured in the vertical axis. Main rotor 4P appears in
the vertical axis.
32
An expanded view of the lateral axis is presented in
Figure 2. From this ploL, main rotor, tail rotor, and engine
frequencies with their associated harmonics appear sharply
defined. This recording position (lateral, tail boom)
provided the best overall view of the helicopters'
characteristic frequencies as compared with the fuselage
measurement position. Of particular interest is the strength
of the main rotor IP vibration. Normally, for a two-bladed
helicopter the 2P would be of greatest amplitude. However,
the Legend IP was the frequency of greatest amplitude, next to
that of engine IP. In a full scale helicopter strong IP
indications generally typify a mass imbalance in the main
rotor system or the result of a blade being "out of track".
During test flights of the Legend no unusual fuselage motion
or blade track anomalies were observed. Therefore, the reason
for the Legend's strong main rotor IP response is unknown.
One possible cause of this relatively high 1P reading could be
the presence of an airframe natural frequency or mode which is
near iP. Particularly suspect in this case would be the
presence of a lateral bending mode. The best means for
resolving this question would be to conduct a vibration survey
of the fuselage (airframe) . Such tests are planned in the
future. As a final note, increasing the speed of the main
rotor system from 1502 (Figure 2) to 1650 RPM (Figure 1)
resulted in a decrease in the level of vibrations as measured
at the tail boom. Such a change in response with change in
33
frequency supports the theory that the high IP response is at
least partially due to the proximity of a fuselage mode to IP.
3. Main Rotor Mass Imbalanck Measurements -.
The effects of a mass imbalance in the main rotor
system were studied by measuring the vibration level at the
tail boom in both the lateral and vertical axes. As noted in
para. 2, this tail boom location was chosen because it
provideO the best response to the Legend's characteristic
frequencies (main rotor, tail rotor, etc.) . Results from this
phase of testing are presented in Appendix E.
Appendix E, Figure 1 shows lateral vibration measurement
results with the main rotor flybar balance mass 1.75 in.
inboard from the normal trim position. This is equivalent to
introducing into the system an unbalance delta of 0.5 in.-lbs.
The characteristic frequencies plotted in Figure 1 show that
a mass imbalance in the main rotor system of this magnitude
results in an increased main rotor induced IP amplitude in the
airframe. As compared with Appendix D, Figure 2, a change of
.04 EU (.04 G) in the airframe lateral response resulted from
main rotor 1P excitation with introduction of .05 in.-lbs. of
unbalance. Main rotor 2P appears unchanged as expected. Also,
engine IP showed it is unchanged, as expected.
Figure 2 of Appendix E is a combined lateral/vertical plot
of tail boom vibration levels with the balancing mass located
3.5 in. inboard from the normal balancing mass trim position.
34
This is equivalent to doubling the unbalance delta from .05
in.-lbs. to .10 in.-lbs. As expected, tail boom lateral
measurements show higher levels of vibration than those
measured in the vertical axis, though main rotor 1P and 2P are
clearly identifiable in both axes. Main rotor 1P shows
significant increase in amplitude over normal levels. Main
rotor 10P showed the greatest increase in amplitude. Small
changes in main rotor operating speed did not appear to change
vibration levels as recorded during these tests, indicating
that if there was proximity with any airframe modes, we were
not dwelling precisely on a mode. Tail rotor 1P shows a
significant increase in amplitude as compared with normal
balanced main rotor operation. This is due to the induced
motion of the tail rotor hub. An expanded view of lateral
axis measurements is presented in Figure 3. With the balance
mass 3.5 in. inboard (for an unbalance of .10 in.-lbs.), tail
rotor 1P has increased to an amplitude of approximately .43
EU (.43 G). This is an increase of approximately .30 EU (.3
G) when compared with normal main rotor operation (Appendix D,
Figure 2). A summary of the results of the tests with rotor
unbalance is given in Table V.
35
TABLE V. MAIN ROTOR MASS UNBALANCE TEST RESULTS
Mass UnbalanceLocation (inboard) 0 1.75 in. 3.5 in. 3.5 in.
Main rotor RPM 1502 1687 1650 1687
IP Lat. Vibrations .134 G .175 G .333 G .432 G
IP Vert. Vibrations .217 G
NOTE: ------ " denotes measurement not recorded.
C. PACIFIC RPV CO. BRUISER GROUND TEST MEASUREMENTS
As stated in Method of Tests, baseline vibration
measurements of the Pacific RPV Co. Bruiser were limited to
ground tests with the accelerometers mounted at the aft skid
support (centerline) position on the fuselage. Appendix F
contains the results of the Bruiser ground tests. AppendixF,
Figures 1 through 3 contain the results of the
lateral/vertical vibration measurements. Figure 1 shows a
distinct main rotor 4P in both the lateral and vertical axes.
At a 4P frequency of 77.5 Hz., the amplitude recorded at the
fuselage was .335 G in the lateral and .674 G in the
vertical axis. Tail Rotor IP is slightly more sharply defined
in the lateral axis as compared with the vertical (Figure 2).
Engine iP may be identified in both axes (Figure 3), though it
is of greater amplitude in the vertical.
Figures 4 through 8 are views of the vertical axis
accelerations up to a maximum frequency of 400 Hz. Figure 4
annotates the main rotor IP, at an amplitude of .103 G. Note
36
that this amplitude for 1P is only 23% of main rotor 2P
(Figure 5) and 13% of main rotor 4P (Figure 6) . Main rotor 4P
was the characteristic frequency of greatest amplitude
recorded during ground tests of the Bruiser. This result
compares favorably to full scale 4-bladed helicopter
vibrations, where the n/rev, operating frequency is typically
of highest value. Tail rotor 1P (Figure 7) and Engine 1P
(Figure 8) were quickly identified through the gearing ratios
(6 * MR 1/rev., and 8 * MR 1/rev., respectively), with
amplitudes within .10 G of each other.
As with the GMP Legend, the vertical axis accelerometer
indicated the highest level of vibration for the fuselage
accelerometer mount. Attempts to measure vibration levels of
the Legend on the ground resulted in only the engine 1P being
identified. This was not the case with the Bruiser, as all
frequencies of interest were identified without the RPH
becoming "light on the skids". The ability to obtain a
characteristic frequency spectrum of the Bruiser while still
on the ground is indicative of the higher level of vibrations
(as compared with the Legend) transferred from the main rotor
system during movement (i.e. rotation) of its dynamic
components to the fuselage.
37
VII. CONCLUSIONS AND RECOMMENDATIONS
A. CONCLUSIONS
1. General
a. The data acquisition system was safe to operate
during all ground and flight tests.
2. The GMP Legend
a. The data acquisition system was of adequate
sensitivity and bandwidth to detect and display the
characteristic frequencies of the two-bladed Legend's dynamic
components.
b. The tail boom, lateral axis, was determined to be
the optimum location for detection of all characteristic
frequencies of the Legend's dynamic components.
c. Using the tail boom-mounted, lateral axis
accelerometer the author was able to identify main rotor, tail
rotor, and engine frequencies and their harmonics as displayed
on the SD380 spectrum analyzer.
d. The Legend exhibited a vibration profile similar
to that of a full scale helicopter with the following
exception: the Legend's engine IP and main rotor iP vibrations
showed amplitudes that were too high when compared with normal
two-bladed helicopter dynamic response. The high main rotor
38
IP response is probably due to the proximity of an airframe
lateral mode to IP, as evidenced by the sensitivity of the IP
response to changes in RPM. For the high engine IP, no
comparison can be made, since the equivalent full-scale
helicopter has a turboshaft engine instead of a reciprocating
engine.
e. The Legend was ideally suited as a remotely
piloted helicopter for both pilot training and
instrumentation development. Instrumentation methods
developed using the Legend were effectively applied to
subsequent Bruiser ground tests.
3. The Pacific RPV Co. Bruiser
a. The data acquisition system was of adequate
sensitivity and bandwidth to detect and display characteristic
frequencies of the four-bladed Bruiser's dynamic components.
b. The fuselage accelerometer location, vertical
axis, was satisfactory for the detection of vibrations
experienced within the Bruiser's airframe.
c. Main rotor, tail rotor, and engine frequencies,
including harmonics, could be distinguished easily from data
collected during ground tests of the Bruiser.
d. The Bruiser exhibited vibration traits similar to
a full scale four-bladed helicopter. Main rotor 4P was the
frequency of greatest vibration amplitude, with engine and
tail rotor of lesser value.
39
e. Similarity of the Bruiser's vibration profile to
that of a full scale helicopter will assist the researcher in
correlating future Bruiser HHC operation test results to that
of full scale four-bladed operation.
B. RECOMENDATIONS
1. General
a. Continue use of the Legend for light weight
instrumentation development and pilot proficiency.
b. Modify the Bruiser with a HHC system that will
operate in an open loop mode.
C. Conduct tests with the HHC-modified Bruiser to
examine the effect of HHC has on helicopter vibration and
performance in hover and forward flight.
2. Specific
a. Conduct further tests of the Bruiser using tail
boom locations for accelerometer mounting.
b. Conduct hover flight tests of the Bruiser to
compare in-flight vibration levels with those experienced
during ground test operations.
c. Develop an instrumentation system which will
telemeter vibration data to a ground receiving station without
a "hard wire" connection.
d. Conduct laboratory vibration surveys for both the
Legend and Bruiser to identify both mode shapes and
frequencies of the primary airframe modes.
40
e. As part of the vibration surveys recommended in
para. d., conduct a detailed survey of the response of the
Legend airframe to excitation from 1450 RPM to 1700 RPM in an
effort to explain the high IP response levels measured during
the model flight test program reported in this thesis.
41
APPENDIX A
low IMPEDANCE, VOtIA(W MOr
OUAR1Z ACCELEROMETERwitl, b,,it In a ... ifilie,
Series 302 A
bt tilt-ill imit ygaini atuipliftier: ettlances resoltitionis highi level (5V), low impedance ( 100 01111)
anlalog output r~(htives lonig coaxial or 2 -wite cables'F
it iverted, isolated comtession struicture
*standaorized sentsitivity: suppressed resonanice
*simlplifie~d systems: low fper cltatel cost ij hsetisitive to cable leigit or motiotiL
Use Model 302A for roultie iteasurenter ofvibratiost anid shock ill laboratory, field. flight, vehicularand itdtmstrial applications.
SPECIFICATIONS: ModelNo. 302AModel 302A. a piecistori rlimatz accelerotttei, Range. FS (+5 volt otltitli g 500
mt ast tres the accolet atiml asptect of shotck mnid vilmtiton Resolottonl g 0 01ttolioll front Ig to 500g. over a wide freeljettcy faiige Useltloverranlge g 1000oatd mit tdmt aiver so eitvir otitenal cotnditiotns Sentsitivity Senisitivity ft 2%) ITtV. g tOismsati~fardizedi a 0mnV g Lieits tat tatttc, Rfesonamt Freqte'cy (itomiitd) lit 30 000
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302At32 tteastires lrattsirttl evetits to005 sicondit Overload Retco~ery Its 10ilttt attolt Output Ittutedanice 011111t 100
Strain Sctsitivily g file] inl 001Qul 12 aCCole 0tottteteS are inlstalled by ClaittPItFig Trallsveise Senitivity % -5the pr ucisiot base Sot laces; itt ittite contlact wilth tilt, lemperatinte Cueltitcient I F 003adjacomt strtitclme ofltite test (bject, usuially by tteatts of lemupeittme Range 'F -1001(,1250Iit tmva stit; bet ylliumt cnlttpfi stuff Adh esive attt ita~tttelic Vibt alirt 'SltotL (max) g 2000 5000
ttttt l~tig bases fatcilitate qick insiallaiomt lot str t ctm tt Weight gtin 25tttimtg Sittce Ite lotce ttovittg file imitmtriit is itats 'Exetaiomt fctmtstalt cutmettll ttA 21020tmitled Iflot lits itloreface. it is imtportailt 1lttI the Voltage to Cituremt Regtulator VDC 124 to27ttalirg -tit face be mtacitited ftat For severe shtock Optionsil Models. Model Noetmvimtmtmnlttltte optotiai solder piit Cnttitolo iltor StKkilc to seomdtt 302AU2has tirved moote reliabtle thtan coaxial romtpovtettts Il-gt Sells 1300m.' g litqt suzel 302803
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itisti in tos tSae offered itt assembtledl kit for fit. as illitus fighq Feeqf IOUi I Ivinst~in of 302A04) 302A00,tt atet. comupilete and toadly to itusall fty coutiuectiq lto txat I inc Citbte 30CAAYom I eatiit illstt utiuteult andl Offer ate Stautditfpi ~tot t Use pt clix ii tlo specify lou tttettc seatling r g 1 1302Atilt
1 t it gt otmtti isoltior. loriget little cotistait. htigher __________
set isitivit y aittwelded Itet tttotic seal A vat icty of hattly ftpical Systerts: K 302 A tt, t.e (tStor Impt ptiwet sigttai corttiiimei s (witi to0 witl ttott gait i At'lw.tt,,.t1 ; GKIWA.11)2A,1, g . ''ii-u smrtle or ttti citattitel crittligmtaentis meet mttttalt -I * ..
2042
A C MOItr. BASIC
BATTERY POWER UNITtrvjtr ioli!" tra"S'loctrs
Model 480B
* powers trantsdrticeirs with buflt itt or attachedamrplifier s oil *, 1 **r xi
* surlrplies conistaiii trrent power over thre ~ ~ '' ~signal lead
* mlonlitols not rial or fitlity Ss tem opeirationt'*internal or external chiarge options
Fel powering tow imprtedance piezxoelectric trairsdttcets itswitl* borilt in or attached arriptlier s anld corilig thinrnr tooscilloscopes. mletels or recordls. especially ill field. lownoise or test stil evatitttioi appicationis
IJew miodelil .1809 teatires 127 volt prower. prnoirer bytiou slarvkrid 9 virll batteries rontiecteri iii series, to extoirdtlie ilyirmiic tanget tfrmany ICP tralisdocers to 10 volts colirir SPECIFICATIONS: Model No. 480BOtlher teatrires mrcliidelair extet al DC powur jack ftr rise with fi aimdoie E xcitation VUC 2 7rrgrtronn ecrzgr'ale battery pack Modlel073805 aid a hat F ictitjiol Cr ittPri ioA 2Irly chlar ge iar t lot Ise with Modiel 4868 Battcery C~liatlge Vlag GiNifrIt Battery irritiVltaeeai)i
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lilt i shotl I I ied), or opi~r circuit fyellowl It also ch ecks Battery lile Ill 110fltite y voltage (1i2 7VJ wheni ir e r ocker switch is depit essedi (rirriet ltrs (Irigir & O1111ttr1 il0t 32ii1, trigilt ('rroiit. are Irrirsed li a slrlelilild prlastic ris-e k(ier monitor VI'S 27withi a ririal trarret Connecters are coaxial mindo t0 32 St Wlt n2ctIjari ki; Opliciiorl BNC corriectors itrodets are availableSie(.1129-x15
Scrryri kits inicluirrrrg Mor(e 4806 Power Uit simily VWeigrl findl batteries) or 15 -sptr'srliig arrotier irig ITrese kits contain tire sensor (fl airs
ii. irsr abe 101lorig). trowui rull, scope cable (3 Ift Opliexnal truodels: Moriel NotIrs ip lt run ti;itirtliii BNCI anI ~accessories To specifyaId 1r lix Witli BlJC it tit and outlii 4802K ti tI ill srisri miodel iii utter arid addi prefix cost to tire sensor Withi gail) Xl1. X 10. X 100 480006pieUreit fiet KR to specify kit with NiCd balrrly a id chatiger Model .180D06 with BNC s 480009oltinr g K11302A Accelerornicter Kit Wheo cornnectedi loa lirgirtirig. velocity ort acceltrariri 4BO0f8ri111ior gi calls uregol un imprin irriertaice. In eqireiy r espirise Optional accessories:is esseriaity that of tile Imarsrlrcer arid fire tr arshrcei sigi ra Ch arget wilthrnee NiCd Irafter is 488Bi% ririlier ampitlifiedl t arterriater Constant cnn rit ecritation Recligealile external battery track 0731305frrrrowr fvilte linear ity antI cale dt sing atrifty ot flei Systeiri To spiscify a comhprglete power kit icluding piower umit 10 hid'
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43
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LIST OF REFERENCES
i. Wood, E.R., Powers, R.W., Cline, J. H., and Hammond, C.E. "On Developing and Flight Testing a Higher HarmonicControl System", Journal of the Ame.ican HelicopterSociety, January, 1985.
2. McKillip, R. M. "Periodic-Model Following for theControl-Configured Helicopter" , Journal of the AmericanHelicopter Society, July, 1991.
3. Wernicke, R. K.., and Drees, J. M., "Second HarmonicControl", Proceedings of the 19th Annual Forum of theAmerican Helicopter Society, May, 1963.
4. Daughaday, H., "Suppression of Transmitted HarmonicRotor Loads by Blade Pitch Control", USAAVLABS TechnicalReport 67-14, November, 1967
5. Balcerak, J. C., and Erickson, J. C., Jr., "Suppressionof Transmitted Harmonic Vertical and Inplane Rotor Loadsby Blade pitch Control", USAAVLABS Technical Report 69-39, July 1969.
6. Shaw, J., "Higher Harmonic Blade Pitch Control forHelicopter Vibration Reduction: A Feasibility Study,"MIT report ASRL TR 150-1, December, 1968.
7. Sissingh, G. J. and Donham, R. E., "Hingeless RotorTheory and Experiment on Vibration Reduction by PeriodicVariation of Conventional Controls", RotoicraftDynamics, NASA SP-352, February, 1974.
8. McHugh, F. J. and Shaw, J., Jr , "Benefits of HigherHarmonic Blade Pitch: Vibration Reduction, Blade-LoadReduction, and Performance Improvement", Proceedings ofthe American Helicopter Society Mideast Region Symposiumon Rotor Technology, Philadelphia, Pa., August, 1976.
9. Hammond, C. E., "Helicopter Vibration Reduction viaHigher Harmonic Control", Proceedings of the RotorcraftVibration Workshop, NASA Ames Research Center, February,1978.
10. Wood, E. R., Higman, J., and Kolar, R., "Higher HarmonicControl Promises Improved Dynamic Interface Operations",AGARD Proceedings of the 78th Flight Mechanics Panel,May, 1991, Seville, Spain.
62
11. Hintze, C. J., "Construction and Use of a RadioControlled Model Helicopter for Research", Master'sThesis, Naval Postgraduate School, Monterey, CA, March1985.
12. Cotten, R. P., "Hover Performance of a Remotely PilotedHelicopter", Master's Thesis, Naval Postgraduate School,Monterey, CA, December, 1986.
13. Scott, J. G., "Establishment of a Remotely PilotedHelicopter Flight Test program for Higher HarmonicControl Research", Master's Thesis, Naval PostgraduateSchool, Monterey, CA, June 1990.
14. Webb, C. D., "Initial Design Study of Existing FlightControl System of RPH and Feasibility Study ofImplementing HHC on the SH-60B," Master's Thesis, NavalPostgraduate School, Monterey, CA, September, 1990.
15. McGovern, J. J., "Flight Operations for Higher HarmonicControl Research", Master's Thesis, Naval PostgraduateSchool, Monterey, CA, March, 1991.
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INITIAL DISTRIBUTION LIST
No. of Copies
1. Defense Technical Information Center 2Cameron StationAlexandria, Virginia 22304-6145
2. Library, Code 52 2Naval Postgraduate SchoolMonterey, California 93943-5000
3. Chairman, Code AA/Co 2Department of Aeronautics and AstronauticsNaval Postgraduate SchoolMonterey, California 93943-5000
4. Professor E. Roberts Wood, Code AA/Wd 2Department of Aeronautics and AstronauticsNaval Postgraduate SchoolMonterey, California 93943-5000
5. Professor Richard M. Howard, Code AA/Ho 1
Department of Aeronautics and AstronauticsNaval Postgraduate SchoolMonterey, California 93943-5000
6. Mr. Kevin M. Ransford 2P. 0. Box 616Lexington Park, Maryland 20653
64